1. Field of the Invention
The present invention relates to airfoils used in a gas turbine airfoil, and more specifically to an airfoil having a platform.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a turbine section includes a plurality of turbine blades and guide nozzles or vanes on which a hot gas stream reacts to drive the turbine. This hot gas stream passes through and around the turbine blades. A hot gas migration phenomenon on the airfoil pressure side is created by a combination of hot flow core axial velocity and static pressure gradient exerting on the surfaces of the airfoil pressure wall and the suction wall of adjacent airfoils. As a result of this hot gas flow phenomena, some of the hot core gas flow from an upper airfoil span is transferred toward a close proximity of the platform and therefore creates a high heat transfer coefficient and high gas temperature region at approximately two-thirds of the blade chord location. FIG. 1 shows a cut-away view of the vortices formation for the hot glow gas migration across the turbine flow passage, and shows a hot spot on the platform of each blade on the pressure side 18.
A Prior Art blade with platform is shown in FIG. 2. The blade includes a root 10, a cooling fluid passage 12, a platform 14, an airfoil 18, and a tip 19. A fillet region 16 is formed between the airfoil and the platform. Cooling of a blade fillet region 16 by means of conventional backside convective cooling method yields inefficient results due to the thickness of the airfoil fillet region 16. On the other hand, drilling film holes 17 at the blade fillet to provide for film cooling produces unacceptable stress by the film cooling holes 17. A line of film cooling holes 17 along the lower section of the blade for cooling the blade fillet region 16 would be located in the region of the airfoil having the highest pull stress levels, thereby providing a point of weakness at the highest stress points on the blade.
U.S. Pat. No. 6,341,939 B1 issued to Lee on Jan. 29, 2002 entitled TANDEM COOLING TURBINE BLADE discloses a turbine blade with a central cooling air passage and a metering hole leading from the central passage and onto the outer surface of the platform around the transition region of the blade for cooling the transition region (space between the airfoil and the platform). However, the Lee invention does not uncouple the airfoil from the platform as does the present invention, among other differences.
U.S. Pat. No. 5,340,278 issued to Magowan on Aug. 23, 1994 entitled ROTOR BLADE WITH INTEGRAL PLATFORM AND A FILLET COOLING PASSAGE discloses a turbine blade with a cooling fluid passage connecting the core passage of the blade with the damper or dead rim cavity for the purpose of cooling the fillet of the platform and airfoil transition. No cooling air passes onto the outer surface of the airfoil platform or airfoil, and the platform is not uncoupled from the airfoil as in the present invention, among other differences.
U.S. Pat. No. 5,382,135 issued to Green on Jan. 17, 1995 entitled ROTOR BLADE WITH COOLED INTEGRAL PLATFORM shows a turbine blade with a platform having a plurality of cooling holes located on the pressure side of the blade for cooling the platform. A row of cooling holes closest to the airfoil surface are supplied with cooling air from the core or central passage of the blade, while an outer row of cooling holes are supplied with cooling air from the dead rim cavity below the platform. The Green invention does not provide for the uncoupling of the platform from the airfoil as in the present invention, among other differences.
One alternate way of cooling the fillet region is by the injection of film cooling air at discrete locations along the airfoil peripheral into the downward hot gas flow to create a film cooling layer for the fillet region 16. However, in order to achieve a high film effectiveness level, the discrete holes used in this type of film cooling injection have to be in a close pack formation. Otherwise, the spacing between the discrete film cooling holes and areas immediately downstream of the spacing are exposed to less cooling or no film cooling air at all. Consequently, these areas are more susceptible to thermal degradation and over temperature. On the other hand, the close pack cooling holes at the blade lower span becomes undesirable and the stress rupture capability of the blade is lower.
An object of the present invention is to reduce or eliminate the high heat transfer coefficient and high gas temperature region as well as high thermal gradient problem associated with a turbine blade platform.
Another object of the present invention is to uncouple the platform from the airfoil of the blade in order to reduce stress from thermal gradients between the two parts of the blade.